THE UNIV E R S I T Y OF MICHIGAN COLLEGE OF ENGINEERING Department of Electrical Engineering Space Physics Research Laboratory SOUNDING ROCKET FLIGHT REPORT NASA 18.01 Thermosphere Probe Experiment Prepared on behalf of the project by D. R. Taeusch and G. R. Carignan ORA Project 07065 under contract with: NATIONAL AERONAUTICS AND SPACE ADMINISTRATION GODDARD SPACE FLIGHT CENTER CONTRACT NO. NAS 5-9113 GREENBELT, MARYLAND administered through: OFFICE OF RESEARCH ADMINISTRATION ANN ARBOR May 1966

TABLE OF CONTENTS Page LIST OF ILLUSTRATIONS v PROJECT PERSONNEL vii 1. INTRODUCTION 1 2. GENERAL FLIGHT INFORMATION 2 3. LAUNCH VEHICLE 3 4. NOSE CONE 8 5. THE THERMOSPHERE PROBE (TP) 10 5.1 Omegatron 10 5.2 Electrostatic Probe (ESP) 23 5.5 Support Measurements and Instrumentation 25 5.3.1 Sun-earth aspect determination system 25 5.3.2 Telemetry 26 5.3.3 Housekeeping monitors 27 6. ENGINEERING RESULTS 28 7. DATA ANALYSIS 29 7.1 Trajectory 29 7.2 Ambient N2 Density 29 7.3 Temperature 33 8. REFERENCES 39

LIST OF ILLUSTRATIONS Table Page I. Omegatron Data 21 Figure 1. Nike-Tomahawk dimensions-6.5-in. diam payload.5 2. Nike-Tomahawk vehicle with payload. 6 3. Nike-Tomahawk vehicle with payload. 7 4. Nose cone dimensions. 9 5. Payload diagram (B-E236). 11 6. Block diagram (B-E273). 12 7. Thermosphere probe in nose cone. 13 8. Expanded view of omegatron system. 14 9. Breakoff configuration. 15 10. Omegatron envelope. 17 11. Magnet. 19 12. Calibration. 22 13. Electrostatic probe feedback network. 23 14. AV generator of ESP. 25 15. Trajectory program output format. 30 16. Trajectory with timing. 31 17. a versus time. 32 18. Peak-BG current versus time. 34 v

LIST OF ILLUSTRATIONS (Concluded) Figure Page 19. Gauge temperature versus flight time. 35 20. Ambient N2 density versus altitude. 36 21. Ambient N2 temperature versus altitude. 37 vi

PROJECT PERSONNEL N. W. Spencer Principal Investigator GSFC G. R. Carignan Project Director Univ. of Mich. R. F. Atmore Thiokol Representative Thiokol Corp. D. J. Beechler Engineer Univ. of Mich. C. F. Bihlmeyer, Jr. Draftsman Univ. of Mich. L. H. Brace Charged Particle Scientist GSFC B. J. Campbell Design Draftsman Univ. of Mich. E. L. Degener Technician Univ. of Mich. P. L. Freed Chief Technician Univ. of Mich. G. K. Grim Engineer Univ. of Mich. D. N. Harpold Calibration Physicist GSFC W. G. Kartlick Omegatron Instrument Maker Univ. of Mich. B. C. Kennedy Omegatron Engineer Univ. of Mich. T. B. Lee Electrostatic Probe Engineer Univ. of Mich. J. C. Maurer Payload Engineer Univ. of Mich. D. L. McCormick Machinist Univ. of Mich. R. L. Navarro Project Engineer Wallops Island H. B. Niemann Neutral Particle Scientist Univ. of Mich. D. T. Pelz Calibration Physicist GSFC Go T. Poole Programmer Univ. of Mich. N. E. Peterson, Jr. Vehicle Manager GSFC G. F. Rupert Telemetry Engineer Univ. of Mich. R. W. Simmons Data Reduction Manager Univ. of Mich. M. D. Street Technician Univ. of Mich. D. R. Taeusch Neutral Particle Scientist Univ. of Mich. G. S. Woodson Programmer Univ. of Mich. vii

1. INTRODUCTION This report describes and discusses the results of the launching of NASA 18.01, a Nike-Tomahawk sounding rocket. The payload was the Thermosphere Probe (TP), described by Spencer, Brace, Carignan, Taeusch, and Niemann (1965). The TP is an instrumented ejectable package developed by this laboratory in cooperation with the Goddard Space Flight Center, Laboratory for Atmospheric and Biological Sciences (GSFC) for the purposes of studying the variability of the earth's atmospheric parameters in the altitude region between 120 and 350 km. The NASA 18.01 payload included an omegatron mass analyzer (Niemann and Kennedy, 1966), an electron temperature probe (Spencer, Brace, and Carignan, 1962), and a sun-earth aspect sensor. This complement of instruments permitted the determination of molecular nitrogen density and temperature, and electron density and temperature in the altitude range of approximately 140 to 300 km. A general description of the payload kinematics, orientation analysis, and data reduction techniques is given by Taeusch, Carignan, Niemann, and Nagy (1965). The orientation analysis and nitrogen data reduction were performed at this laboratory and the results are included in this report with a discussion of problem areas and probable errors. The electron temperature probe data were reduced at GSFC, and are not discussed in this report. The payload described herein was launched 12 hours prior to a similar one (NASA 6.11) described in a separate report. The purpose of this dual launching was to establish the diurnal variation of the parameters measured to provide extra meaning to their use in studying the effect of the energy input to the atmosphere. 1

2. GENERAL FLIGHT INFORMATION The general flight information for NASA 18.01 is tabulated below. The geophysical indices listed, the 1.0.7 cm solar radio flux, F10 7; the five monthly averages of the solar 10.7 cm flux preceding the launch, F107; and the geomagnetic index, ap, were obtained from the April, 1965, and May, 1965, issues of "Solar Geophysical Data" published by the U. S. Bureau of Standards F10.7 is given for the day preceding the launch and ap is given for six hours previous to launch for convenient reference to the Harris and Priester (1964) model atmosphere. The Table of Events gives flight times and altitudes of significant events occurring during the flight. Some of these were estimated and are so marked. The others were obtained from the telemetry records and radar trajectory, where applicable. Launch Date: March 19, 1965 Launch Time~ 1309 EST; 1809 GMT Locations Wallops Island, Virginia Longitude~ 75. 04W Latitude 37 54~N Apogee Parameters: Ait-tude 3524o8 km Horizontal Velocity: 569.0 m/sec Flight Time~ 282.5 sec &,_eophysicai Indices~ -lo. 74~3 TP Motion 07 = 75.3 Tumble Period: 1,7548 sec/tumble a 6 Roll Period: ~ 2

TABLE OF EVENTS Event Flight Time Altitude Remarks Event ( Remarks (sec) (km) Lift Off 0 0 1st Stage Burn Out 3.5 (est) 1.4 (est) 2nd Stage Ignition 11.45 6.35 (1,904 fps) 2nd Stage Burn Out 20.75 19.6 (8,384 fps) Despin 43.0 (est) 60.0 (est). TP Ejection 45.0 64.2 Omegatron Breakoff 75.0 (est) 126.0 (est) Omegatron Filaments On-to Mass 28 76.8 128.3 Omegatron to Mass 32 127.5 219.2 Omegatron to Mass 28 167.2 266.6 Peak Altitude 282.5 324.8 LoO.S. 517.0 80.0 5

5. LAUNCH VEHICLE The NASA 18.01 launch vehicle was a two-stage Nike-Tomahawk combination. The first stage was the solid propellant Nike booster, which has an average thrust of 49,000 lb and burns for approximately 5.5 sec. The Nike is 155 in. long, 1.6.5 in. in diameter, and weighed 1550 lb unburned. The center of gravity (CG) was 75.7 in. from the nozzle exit plane (NEP). The second stage was Thiokol's Tomahawk solid propellant motor. The average thrust is approximately 11,000 lb and it burns for about 9 sec. The Tomahawk is 142 in. long, 9 in. in diameter, and weighed 554 lb unburned. The CG was 72.125 in. from the NEP. The payload was 76.9 in. long and weighed 125 lb making the total vehicle 362.5 ino long and weighing 1986 lb. Drawings and photographs of the vehicle are given in Figures 1, 2, and 3. The predicted performance for the vehicle was 3522 km peak altitude at 275 sec flight time. The actual performance, as discussed in detail in a later section, was 324.8 km peak altitude at 282 sec of flight time. 4

= —------------- ~~367. 7 14 5.2 -- 14 1.7 5 __36,6__ 59.936 t 16.5DIA./ 9 D A. FIRST STAGE' SECOND STAGE PAYLOAD viu^~ N NIKE BOOSTER TOMAHAWK ORDNANCE ITEMS ORDNANC*E I TE~M~S /FIRING ~ DESPIN UNIT Q NOSE CONE OPENING PRIMERS. ( BREAKOFF LINEAR ACTUATORS ) DESPIN INITIATION PRIMERS ( SECOND STAGE IGNITER O NIKE BOOS TER IGNITER Figure 1. Nike-Tomahawk dimensions-6.5-in. diam payload.

Figure 2. Nike-Tomahawk vehicle with payloa 6

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4. NOSE CONE A picture of the total payload including nose cone, despin mechanism, and adapter sections is shown in Figure 4. Section A contains the batteries, timer, and pyrotechnics for the opening of the nose cone. Section B contains the plunger and the volume which hoaid the TP (E). Section C houses the ejection spring, plunger piston, and the lanyard negator motor. Section D is the despin mechanism. Dimensions and weights of the system are given on the schematic. The payload is programmed to despin at about 70 km altitude and the TP is ejected and tumbled at 75 km. The breakoff device is removed at about 110 km and the omegatron turned of a new seconds later. The timing for this particular payload was described previously. 8

k&E 10-SS14 7-13. -------— 76.2i5 —-- -5 9'DIA. 1- 262.88 - -~ -.50 DIA. -. —a.aS) — 5./ —-23./2 D | C BI E / 1A C.G. 40. TOTAL WEIGHT = 115 LBS. MOMENT OF INEE IIA ABOUT C.G. - /Z.7 SL _-ft_'.L. H. R.HH.. R ^.... I...T... PART NO. NAME NO. RtO'D DASH NO. MATERIAL DESIGNED BY APROVLED *Y SPACE PHYSICS RESEARCH LABORATORY.AWN *Y.J. C. SCA./5 = /.000 DEPARTMENT OF ELECTRICAL ENGINEERING CHECKED.Y DAT / G THE UNIVERSITY OF MICHIGAN D I M E NS 10 NA L S PE CS. ANN ARBOR, MICHIGAN FOR TOMA/HA-WK 12.05 I ____________ T.P. EJECTION NOSE CONE PROJECT UNLESS OTHERWISE SPECIFIED TOLETANCES ARE: DIM. ENDING.00.030 ANGULAR DIM. DWG. NO. B- I 4 - 3 I ENDIN..00 ~.010 ~ 30 McIN. N - n 14 - 3 Figure 4. Nose cone dimensions.

5. THE THERMOSPHERE PROBE (TP) The TP used for the NASA 18.01 payload was a cylinder 32-7/64 in. long and 6 ino in diameter weighing 43 lb. The prime instruments for this payload were an omegatron mass analyzer (Niemann and Kennedy, 1966), and an electron temperature probe which utilized two electronic systems. Supporting instrumentation included an Adcole sun aspect sensor and earth sensor for the determination of the TP aspect. The diagram shown in Figure 5, shows the instrumentation and supporting electronics location; and Figure 6 shows the block diagram. Figure 7 is a picture of the completely assembled TP. 501 OMEGATRON The omegatron used in this payload was of the type described by Niemann and Kennedy (1966). An expanded view of the system is shown in Figure 8. Table I lists the operating parameters of the gauge and associated electronics. The characteristics of the linear electrometer amplifier current detector, used to monitor the omegatron output current, are also listed. This omegatron was the first flight model to utilize a ceramic breakoff device which allowed vacuum sealing of the gauge. The breakoff configuration is shown in Figure 9. A new omegatron envelope was designed for this breakoff device and is shown in Figure 10. The magnet used for this system is shown in Figure 11. The calibration of this gauge, and that one used in NASA 6.11, was per-.formed in Februaryo 1965, at GSFC. The system used was an oil diffusion pump ca.libration system ndr er e ervision cf Mr, Carl Reber of GSFC. Other gauges used for reference were (1) a double focusing 1L80~ magnetic deflection spectrometer, (2) a Westinghouse Bayard-Alpert (BA) gauge) (3) two Veeco BA gauges, calibrated by Ball. Brothers, and (4) an AW5966 BA gauge. The two 7 —co gauges were used as the standard for this calibration. They had been calibrated aga.inst a McLeod gauge to a stated absolute certainty of better tha.n + 2;o%. As stated previously, the NASA 6.1.1 omegatron was calibrated on this system at the same time, therefore ~the relative accuracy between the two omegatrons is believed better than + SlOo. A final relative calibration was perfcormed at SPRL on February 26, 1965, at which time the NASA 18.01 omegatron was refocusedO The NASA 6.11 calibration was used as the standard after the refocusing and a new sensitivity was determined for the NASA 18.01 gauge. The final calibration is shown in Figure 12. 10

K&E V1911 111-4l3S 322 144 13 2 OM BREAK-OFF -ELECTRONICS FOR SENSOR (NOT USED IN 6.11) EARTH CELL (NIKE- TOMAHAWK) / ADCOLE ASPECT SENSOR / ELUNAR ASPECT SENSOR / (NOT USED IN 6.11) (NASA 6.11)________________ 9 1? 3 1 _ 13 1 1 Hi J _ _ _ _ --'J1 I4 — 216 TRANSMITTER AND BATTERY DECK MEGATRON GAGE / / / / / / TP DE OMEGATRON ELECTROMETER / / / / /6(NOT FOR 6.11) / a DER / / / / / ETP DECK 1 FILAMENT fREGULAtt DECK / / SCO DECK -OSCILLATOR DECK OMx5 AMPLIFIER DECK /COMMUTATOR DECK LOGIC DECK FLIGHT CONTROL DECK ENGINEER JM DRAFTSMAN CB SPACE PHYSICS RESEARCH LABORATORY DECK LAYOUT DEPARTMENT OF ELECTRICAL ENGINEERING NASA 6.11 &12.05 2 —6 UNIVERSITY OF MICHIGAN 1 -18-65 ANN ARBOR, MICHIGAN B E236DATE C2 Figure 5. Payload diagram (B-E236).

PULL OFFS po2 po3 po4 po1 OMEGATRON SYSTEM VECTOR SCO's amplifier -CALISAWT- 70 kc -QM ledexy^^~~ oscillotor trol filament regulator 40kc OM gage CAB & THEM CONTROL LEDEX / tLl!voltage monitors O F C MT ^ ^ FILAMENT' ---— FILAMENT COMMUTATOR amp control / TIMER 1 sectio' 7.35kc COMM 30 segment lcontroL ~! filament control // / _ --- ex., ---— C^ 5 ^ -----------------— AE -T TR_ ^ ^1XSTEIPM PHYSICS- 22Rkc ELSP INTERNAL " POWER I l_9-HR-l$'S J ADCOLE SOLAR ASPECT -- 10.5kc ASPECT SYSTEM ENGINEER J M DRAFTSMAN CB SPACE PHYSICS RESEARCH LABORATORY BLOCK DIAGRAM 2-2&6 DEPARTMENT OF ELECTRICAL ENGINEERING T M SFR SH RE PROB 2UNIVERSITY OF MICHIGAN 12-12ANN ARBOR, MICHIGAN B-:E273 Do ^ C-1A Figure 6. Block diagram (B-E273).

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16 FREOQUENCY VCO AMPLITUDE 28 &.32 FREQUENCY 32 \A^'X\'\ ^\. \ ^^ ^^'s^'16 AMPLITUDEA I BEAM ADJUST MAGNET\.^l ^ ^OFA FLNG DECKS DECK REGULATOR AUXILIARY D EC K DECK M j EGAT RONA AM PLIFI~ER OM EGAT RONN ADAPTER OMEGATRON SYSTEM - EXPANDED BREAK ~ OFF CONFLAT FLANGE FM-s Figure 8. Expanded, view of omegatron system.

r^\^ ~~A-OZO-075 V"0/ B~~~3REAK -OFF? CUA'RO O3- ^-0o2 APPROQX. LOC. OF IIG ECQ. 5PACED SPOT WELMP Z/ OyPE'RATION -PROCETTDUIRE:_ _ ____ _ ________________ I. MACHINE: C-EO-0?-BTEIAK-OF-FBASE~ WITHOU. _ —CON P-L AT G~ROOVE ~ —------------------------------- J _ __ __. _-45 __BREAK-OFF CE^MIC T^C^ T co. E. MACHINE: C-020-077 —BREAW —0rP-f HAT J HEU-A-RC. WEL7-P sloe5-lco 5_DTIIN.TMO ___T^A MOID IF IEID C:RA\,oRTl S\W\AG~E:LCO< *Cmo00-1-4^.W-51( TO HAT. _ g —4-31gS CTAFRD nlNGC V >ELK fo: LEAK. CRETCkl- HELI-A'RC \W^ELTD - M^U5>T -BE: VACUUM^ 71S~H-r. ~~ ^o0 NARAT.\OR ______ 3. 5ENTD T>E.TAILS C-/20-074- 4- C:"020-OTT TO COOT^. ~POT^CELIN CO. TO -BRAZE CQORS':BRF-AV,- QEF CERAMIC SETAL -RING~ -TQ — --- -- ---------------:.~TAILS C-02O-0T4- 2 C Qg -(D7. — - --------------------- 4.- 5POT WEL7D A-020-OT5 — BRZEAYC- OPFP GUARID TO C-020Q^- -— 07-7o-^^rr -AP —------ 7BREAY,-P0F9- HATr AT U. OFM. AFTER -5RAZI MGi QP CERAM IC SEAL.. AJI A-020Q7;'OT~ _BEK-FPCUKJ____ 5. LEAYI CHF-Ckl COMPLETE: UN17. IC-Og 4 B-OO? Ge. RNAL M^ACHININGi 01 CONPLAT SiROOVE: ON C-020O-OT4- - 4__4-020 3_QEGTQN*3E7-F SP^Ic'BREAk W- 0=FT -BA'SE:, I C-020-077 gR~lQPhA ________' C-Q0aEA0. OFF <~ CUA"A' -—' —~~~~~~~~~~~~~~~~~~~~~~~~~~~~L H. R~ O. HO. O -hiE"Q-"! PARTNO. 5AME 5]=C EP". NO. REQ-D. _________D S NO _____________ ________ MA E I L ________ SPACE PHYSICS RE!R ABRAOR IC DEPARTMENT OF ELETALNG NG "g.N. 1~L3 " ~.C THE UNIVERS O: G_ 1 3~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~I. SN ETDN 0 00 ANGLAR DIM.-74,F..NO. C-2O-"? TO0 C-~::~?ORCEL ____________CO.________TO_____________________:________C_______'_____________-______________CE_____ A M IC_ DM.EDIG 00SE0AL _______ 0 I. ____N G % O >/ - <^ Figure 9. Breakoff coirfiguration.~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~I [DE-TAILS C-O80- O& li~ C-O20- OF1?

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TABLE I OMEGATRON DATA Omegatron Gauge Parameters Beam current 4.05 pa Electron collector bias +77.5 v Filament bias -87.1 v Cage bias -0.205 v Top bias -0.405 v RF amplitude M28 2.69 v P-P M32 2.67 v P-P RF frequency M28 123.81 kc M32 109.16 kc Monitor Filament OFF 3.857 v ON 4.547 v Beam OFF 3.635 v ON 2.298 v Thermistor pressure* 1.606 v Bias 4.058 v RF M28 3.942 v M32 2.605 v Calibration Sensitivity 4 x 10-5 amp/torr Maximum linear pressure () 6 x 1 torr Electrometer Amplifier Range Range Indicator Range Resistor ZPV 1 0.0 v 3.16 x 1010 5.00 2 ~ 0.7 v 1.0 x 1011 5.00 11 3 1.4 v 3.16 x 101 5.00 4 - 2.1 v 1.0 x 1012 5.00 5 2.8 v 3.16 x 1012 5.01 6 3.5 v 1.0 x 1013 5.04 Calibrate 0.819 v Miscellaneous +28 power current all on 300 ma Preflight gauge pressure Magnetic field strength ~ 2200 gauss *Filament off, zero pressure. 21

z U U0~~~~~~~ /^ ~~~~~~~~~~~SENSITIVITY ~ 7869 x 10 AP cc 0 r'1') ~- ^ ~-~~~~~~~~~~~~~~~~~~~~* - INCREASING PRESSURE 0 O X u DECREASING PRESSURE 10 0 I 0 - 13 I0 10 10F~~ ~ NI LET ( 2 PAR^ I 12 IT/CS ) -14 107 10' I0 I00~10 102 101I I101 Figure 12. Calibration.

5.2 ELECTROSTATIC PROBE (ESP) The ESP, described by Brace (1963), consists of a cylindrical probe, which is placed in a plasma and collects current; and an electronic unit, which measures the collected current. The electronic unit consists of a power convertor, AV generator, two three-range current detectors, and relays and associated logic circuits. In this flight the ESP experiment had two objectives: One, a measurement of the probe current, which is normally carried out in TP experiments; and the other, a flight comparison of two current detectors, Detector No. 1 (conventional) and Detector No. 2 (newly developed). Detector No. 1 is of the type which has been employed almost exclusively in the past for these measurements. It uses a ring modulator to obtain a direct current amplification with a floated input configuration. This relatively simple arrangement has certain drawbacks including a tedious process of diode matching, nonlinearity at low currents (Brace 1959), and temperature dependence of the modulator. These drawbacks coupled with "state-of-the-art" advances suggested use of a different technique and the second detector was flown for this purpose. Detector No. 2 employs a difference amplifier obtained through the use of an "off-the-shelf" operational amplifier and a feedback network. Figure 13 shows the network arrangement. With this, the simple relation for probe current, ip e eo/Rg, is obtained for R1 < Rg. TO PROBE ip RI R mR PHILBRIC 0o () AV GEN R2 ip n eo FOR RI < < R2 R2 Figure 13. Electrostatic probe feedback network. 23

The temperature dependence of the detector is minimized by using matched, low temperature coefficient resistances for the feedback network. Consequently, no temperature compensation of the detector was required for the temperature range encountered during these rocket flights. Also by eliminating the tedious diode balancing, temperature test requirements and fabrication of the amplifier, the construction time for the Detector No. 2 was cut nearly 90% from the previous types. The following is a comparison of electrical characteristics of the two detectors: Detector Parameters No. 1 No. 2 Output resistance >. K > 0.2 K Input resistance (for.4 ia) >40 K > 1.0 K Rise time 2 msec 0.5 msec Fall time 5 msec 0.5 msec Lineariy Poor at low Good throughLinearity output out Common mode output (for.4 pa) 50 mV (null) >20.0 mV The result of' comparison in actual probe current measurements proved the Detector No. 2 is better in many ways, i.e., no nonlinearity problem at low outputs, fast rise, and fall times and low insertion loss. Due to the improvement of the frequency response of Detector No. 2, the frequency of the AV can be increased in future flights. This means more samiples for a given time are obtainableo Another improvement in the ESP instrument was a new AV generator. The new TV generator uses an integrating circuit with an operational amplifier,,Fgure 1Lo By selecting a proper integrating capacitor, the temperature compensation procedure is eliminatedo The AV generator output is expressed by: E) - E2 e CR t + E2 Therefore a change of the AV slope is obtained by merely changing the relation of (Ei - E2). These improvements cut down construction time of the AV generator by nearly 80%. 24

RI C EE2 P65 R 2 R PHILBRICK ----- E, 3 (E2-E1) AV SLOPE (E2E CR Figure 14. AV generator of ESP. The following are the specifications of the ESP system: (a) AV slope (dV/dt): HI - 38.5 v/sec LO - 13.1 v/sec (b) Current ranges: Range No. 1 - 20.0 AA full scale (4 v) Range No. 2 - 2.0 pA full scale (4 v) Range No. 3 - 0.4 jA full scale (4 v) (c) Sequence: Range Nos. 1, 2, and 3 2 measurements with Detector No. 1 1 measurement-with Detector No. 2 System calibration while detector is not engaged for probe current measurement 5.3 SUPPORT MEASUREMENTS AND INSTRUMENTATION 5.3.1 Sun-Earth Aspect Determination System The NASA 18.01 TP utilized a sun-earth aspect system for the deterniination of the angular momentum vector of the tumbling TP. This system, designed especially for the TP configuration, consisted of an Adcole Model 2302 shift register, three Adcole Model 1307 solar aspect sensors, and one Adcole Model 1601 earth sensor. The sun sensors were mounted ill the center 25

section of the TP such that the command eyes viewed a plane perpendicular to the long axis of the TP cylinder. The view angle in this plane was 360~, and the resolution was 1.4~o The earth sensor was mounted at one end of the TP cylinder and viewed along the long axis of the TP with a view angle of 1~9'. The particulars of the data reduction from this system are described by Taeusch. et a!o (1965). The solar aspect sensors worked as expected during the flight and aspect was determined to an accuracy of approximately + 5~. The TP angle of attack was determined by the velocity vector reference technique as described in Taeusch, et al. (1965). The earth sensor apparently lost its lens early in the flight and, therefore, no useful data were obtained from it. 5o3.2 Telemetry The payload data were transmitted in real time by a six-channel PAM/FM/FM telemetry system at 240,2 m hz with a nominal output of 2.5 watts. The telemetry system used six subcarrier channels, assigned as outlined below. Transmitter: Driver TRPT-250 Serial Number 1045 Power Amplifier TRFP-2V-1 Serial Number 211 Mixer Amplifier-Type TA48 Serial Number 925 Subcarrier Channels (SCO-type TS58) Nominal iRiG Serial CenterNominal Band Number Frequency Frequency rFunction Response 30 hz PAM 11 1868 7.35 K hz 110 hz 5 data Solar aspect 12 1981 1.0 5 K hz 160 hz and earth cell data'15 1699 14.5 K hz 220 hz Electrostatic i-69 14..... K hz 220 hz probe data Electrostatic 14 1713 22.0 K hz 330 hz rota probe data 16 1733 40.0 K hz 600 hz Omgatron output data Omegatron X5 18 1731 70 0 K hz 1050 hz amplifier output 26

Instrumentation power requirements totaled approximately 30 watts, which was supplied by a Yardney HR-1 Silvercell battery pack of nominal 28.5 v output. 5.3.3 Housekeeping Monitors Outputs from various monitors throughout the instrumentation provide information bearing on the operations of the electronics components during flight. These outputs were fed to a thirty-segmentcommutator which ran at one rps. The commutator assignments are as follows: 1. 0 v calibrate 2. 1 v calibrate 3. 2 v calibrate 4. 3 v calibrate 5. 4 v calibrate 6. 5 v calibrate 7. 5 v calibrate 8. 5 v calibrate 9. Omegatron amplifier range 10. Omegatron amplifier output 11. Omegatron filament voltage monitor 12. Omegatron beam current monitor 13. Omegatron bias voltage converter monitor 14. Omegatron RF voltage and frequency monitor 15. ESP AV monitor 16. ESP AV monitor 17. ESP AV monitor 18. ESP range 19. ESP system (1 or 2) 20. ESProbe (1 or 2) 21. HI or LO AV 22. ESP guard (on or off) 23. Thermistor-filament regulation temperature 24. Thermistor-omegatron gauge temperature 25. Thermistor- omegatron amplifier temperature 26. Thermistor-omegatron gauge internal pressure 27. Thermistor-transmitter temperature 28. 4.5 v calibrate 29. 4.0 v calibrate 30. Battery voltage monitor (1 v out - 6.1 v battery) 27

6. ENGINEERING RESULTS The first flight of the Thermosphere Probe on a Nike-Tomahawk vehicle is considered successful. The demonstrated ability of this combination makes possible plans for launchings at remote locations, and for multiple launchings. The reduced costs and preparation effort should also contribute significantly to the future results of the TP program. The successful application of the "breakoff" device, in addition to increasing the altitude and the accuracy of the N2 measurements, also greatly reduces the complexity of field operations and eliminates the need for much bulky support equipment. The base pressure achieved in the pinched-off gauges on this flight was higher than desirable, but as a result of the experience gained, much lower pressures are confidently expected in future applications. Despin of the vehicle, deemed necessary for a good ejection, was accomplished with a newly designed "yo-yo" system. The preparation for flight of this system brought to light certain weaknesses which will be corrected in future flights. The newly designed electrostatic probe circuitry functioned well in flight. Based on this successful flight and the obvious other advantages, a change ofver to the new circuitry on all future flights is anticipated. The first application of the new sun aspect sensor, especially designed for the TP application, was successful and provides a factor of two improvement in angle-of-attack determination for roughly the same cost as the previous sensor. The earth telescope did not function properly; its failure is attributed to either a lens failure or unknown causeso This problem is being studiedo 28

7. DATA ANALYSIS The telemetered data were recorded on magnetic tapes at two stations, Wallops Island Main Base Telemetry Station and Goddard Space Flight Center Station "A." One set of real time "paper" records for quick look evaluation of the results were also obtained. The omegatron, housekeeping and aspect data were reduced to engineering parameters from paper records, run at 10 in./sec, using a Gerber GDDRS data reader and scanner. The paper records used for data reduction were recorded from the magnetic tape masters. Tracking data for trajectory determination were obtained from FPS-16 and FPW-6 radars. Continuous data were obtained from +10 to +121 sec by FPS-16 radar and from +40 to +318 sec by FPQ-6 radar. 7.1 TRAJECTORY The trajectory and velocity information used to determine the aspect, density, and temperature data as a function of altitude were obtained by fitting a smooth theoretical trajectory to the measured radar data. The theoretical trajectory is programmed for computer solution similar to that described by Parker (1962). The output format is shown in Figure 15. The trajectory is shown in Figure 16. The analysis of minimum angle of attack (a) as described by Taeusch t al. (1965) is also incorporated in the program and the output of the computer furnishes a and cos a versus time, altitude, etc. A plot of 18.01 a versus altitude is given in Figure 17. 7.2 AMBIENT N2 DENSITY The neutral molecular nitrogen density was determined from the measured gauge partial pressure as described by Spencer, et al. (1965), using the basic relationship:'Ani Ui na = _^ a /2 V cos a/ N2 where: na = Ambient N2 number density Ani = Maximum minus minimum gauge number density during one tumble Ui = /2kTi/m most probable velocity of particle inside gauge 29

SPACE PHYSICS RESEARCH LABORATORY UNIVERSITY OF MICHIGAN THURSDAY, SEPTEMBER 23. 1965 NASA 12.05 TP8 LAUNCH TIME (GMT) YEAR 1965 DAY 78 HOUR 18 MINUTE 9 SECOND.000 INITIAL CONDITIONS TIME 98.000 SECONDS FROM LAUNCH ALTITUDE 573530.0 FT RANGE 167690.0 FT VELCCITY 5730.0 FT/SEC FLIGHT PATH ANGLE 70.2000 DEGREES UP FROM LOCAL HORIZONTAL PLANE AZIMUTH 131.2400 DEGREES EAST OF LOCAL NORTH LONGITUDE -75.0372 DEGREES (+EAST) LATITUDE 37.5434 DEGREES (+NORTH) THE "CORRECT' MOMENTUM VECTOR, IS -.66966.67655.30563 PEAK PARAMETERS TIME ALTITUDE RANGE POSITION INERTIAL VELOCITY VELOCITY WRT EARTH FIXED VEL AZIMUTH ELEVATION ALPHA DEGREES FEET METERS FEET METERS FEET DEGREES DEGREES SEC FEET FEET LATITUDE TOTAL TOTAL TOTAL TOTAL X INERTIAL INERTIAL DEGREES METERS METERS LONGITUDE HORIZONTAL HORIZONTAL HORIZONTAL HORIZONTAL_ Y WRT EARTH WRT EARTH RADIANS GEOPOT METE VERTICAL VERTICAL VERTICAL VERTICAL Z COSINE 282.53 1065488 502951 36.92 2925 892 1868 569.1359 116.46.00 73.392 324761 153299 __ -74. 9 _2925 892 1868 569 -1280. 134.25._ 00 1.281 309008 -0 -0 -0 -0 -44.2858 Figure 15. Trajectory program output format.

KM-X 1- 153 1 -5. I 1MILLION FEET (190 MILIES 305 KM 0 0 UJ 6 2.000 FEET.100OKM 16000 PT/SEC ~| 1 2 830<00 FILET E7C0KM OOFSPACE PH I /\00 0 FT/SECS zl~' v | (10 SCCOND INTERVALS) TP ELECTONIC C TIMER N.2 ENGINEER DRAFTSMAN JRP SPACE PHYSICS RESEARCH LABORAT )RY NIKE -TOMAHAWK DEPARTMENT OF ELECTRICAL ENGINEERING TIME PROFILE NASA 1205 65 UNIVERSITY OF MICHIGAN 2-16-65 ANN ARBOR, MICHIGAN BE264DATE C-3A Figure 16. Trajectory with timing.

80 70 u, w 60' 50 - 0( 40 - / downle|upleg 30 - 20 upleg O ALTITUDE (K M)..I I I I I 100 150 200 250 300 Figure 17. ca versus time. 52

Ti = gauge wall temperature V = Vehicle velocity with respect to the earth a = Minimum angle of attack for one tumble AIi, the difference between the maximum (peak) omegatron gauge current and the minimum (background) gauge current is shown versus flight time in Figure 1.8 The background current is also shown in the figure. The background current is the result of the outgassing of the gauge walls and the inside density due to atmospheric particles which have high enough energy to overtake the TP and enter the gauge. The outgassing component is assumed constant for one tumble and effects both the peak reading and the background reading; and, therefore, does not effect the difference. From calibration data obtained by standard techniques, the inside number density, ANi, is computed for the measured current. As described by Spencer, Taeusch, and Carignan (1965). the uncertainty in these data is believed to be + 10.2% rms relative to other measurements using the same calibration system and + 25.1% rms absolute. Ui, the most probably thermal speed of the particles inside the gauge, is computed using the measured gauge wall temperature shown in Figure 19. The uncertainty in this measurement is believed to be + 2.2% rms absolute. V, the vehicle velocity with respect to the earth; is believed known to better than + 1% absolute. It is obtained from the trajectory curve fitting described previously and is the most accurately known quantity obtained from the analysis. Cos a is obtained from the aspect analysis described by Taeusch, et al. (1965). Since the uncertainty in cos a depends upon a, for any given error in a, each particular case and altitude range must be considered separately. As can be seen in Figure 17, the upleg data were obtained for angles of attack less than 30~, which results in an uncertanty in cos a approximately + 3% for an uncertainty in a of approximately + 5~. The upleg data were used as control data since the downleg angles of attack were between 55~ to 35~. The resulting ambient N2 number density, obtained from the measured quantities described above, is shown in Figure 20. The uncertainty in the ambient density due to the combined uncertainties in the measured quantities is + 10.9% rms relative and + 25.4% rms absolute. 7.3 TEMPERATURE The ambient N2 temperature profile, shown in Figure 21, was obtained by 33

-9 10d9 i NASA 12.05 WALLOPS ISLAND, VIRGINIA MARCH 19, 1965 - 13:09EST 10 O PEAK MINUS BACKGROUND CURRENT ta. 10 -- u. <I. a. Z.:I-.~~~~~~~~',. 0 10 ~ u. 0 I.-o ~~~~~~~.. ~:,. 613 1...0....'. ~ BACKGROUND CURRENT FLIGHT TIME (SECONDS) 10 I I I I 100 200 300 400 500 Figure 1. Peak-BG current versus time. 14

290 289 w 288- - a.. LU 285 I I I II 100 150 200 250 300 350 400 450 500 Fire 19. Gauge temperature versus flight time. 287- c 286_ FLIGHT TIME (SEC. FROM LAUNCH) 100 150 200 250 300 350 400 450 500 Figure 19. Gauge temperature versus flight time.

109 I 0 U I, z \ ALTITUDE (KM) \ 120 150 200 250 300 330 Figure 20. Ambient N2 density versus altitude. 0z\ ALTITUDE (KM) 36

300 - W/'O 200 w TEMPERATURE (-K) 650 700 750 800 850 900 Figure 21. Ambient N2 temperature versus altitude. 37

integrating the density profile to obtain the pressure and then relating the known density and pressure to the temperature through the ideal gas law. The assumption that the gas is in hydrostatic equilibrium and behaves as an ideal gas is implicit. Since the temperature depends only upon the shape of the density profile and not its magnitude, it is believed that the uncertainty in its magnitude is + 5% absolute. NASA 18.01 March 19, 1965 18:09:00 GMT Wallops Island, Virginia Altitude, Density, Temperature, km part/cc ~K 173 6.83 x 109 691 175 6.22 694 180 4.92 701 185 3.90 708 190 3.10 716 195 2.48 723 200 1.98 730 205 1.58 737 210 1o27 743 215 1.02 x 109 750 220 8.29 x 10 756 225 6.70 763 250 5.42 769 235 4.39 775 240o 555 781 245 2.88 786 250 2.36 792 255 1.93 797 260 1.6o 802 265 1.32 806 270 1.09 x 108 810 275 8.99 x 107 814 280 7.43 817 285 6.18 820 290 5.1 823 295 4.26 824 300 3.55 826 305 2.95 826 310 2.46 827 5153 2.21 x 107 827 38

8, REFERENCES Brace, L. H., " Transistorized Circuits for Use in Space Research Instrumentation," University of Michigan, Space Physics Research Laboratory, Report No. 02816 1-3-S, October 1959. Brace, L. H., Spencer, N. W., and Carignan, G. R., "Ionosphere Electron Temperature Measurements and Their Implications," J. Geophys. Res., 68, 5397-5412 (1.963). Harris, I., and Priester, W., "The Upper Atmosphere in the Range from 120 to 800 Km," Goddard Space Flight Center, NASA, Institute for Space Studies Report, 1964. Niemann, H. B., and Kennedy, B. C., "An Omegatron Mass Spectrometer for Partial Pressure Measurements in the Upper Atmosphere," submitted for publication to Rev. Sci. Instr., 1966. Parker, L. T., Jr., "A Mass Point Trajectory Program for the DCD 1604 Computer," Tech. Doc. Report AFSWC-TDR-62-49, Air Force Spec. Weapons Center, Kirtland AF Base, New Mexico, August 1962. Spencer, N. W., Brace, L. H., and Carignan, G. R., "Electron Temperature Evidence for Nonthermal Equilibrium in the Ionosphere," J. Geophys. Res., 67, 151-175 (1962). Spencer, N. WV, Brace, L. H., Carignan, G. Ro, Taeusch, D. R., and Niemann, H. B., "Electron and Molecular Nitrogen Temperature and Density in the Thermosphere," J. Geophys. Res., 70, 2665-2698 (1965). Spencer, N. W., Taeusch, D. R., and Carignan, G. R., "N2 Temperature and Density Data for the 150 to 200 Km Region and Their Implications," Goddard Space Flight Center, Report X-620-66-5, December 1965. Taeusch, D. R., Carignan, G. R., Niemann, H. B., and Nagy, A, F., "The Thermosphere Probe," University of Michigan, Scientific Report 07065-1-S, March 1965. 59

UNIVERSIOF HIGAN 3 9O 03527 1686